One concern in the operation of gas turbine engines is the catastrophic failure of any part of the turbine wheel including the hub, the blades, or both. Because such failures typically occur when the gas turbine is in operation, and because the rates of revolution of turbine wheels in gas turbines are quite high, the resultant high angular velocity translates to high centrifugal forces acting on the turbine wheel. Should there be a failure where any part or all of the turbine wheel cracks, breaks or begins to disintegrate during turbine operation, the presence of this high centrifugal force will cause the separated component to move radially outwardly at high velocity and with substantial kinetic energy.
To prevent damage to surrounding instrumentalities, it has long been common to provide turbines with so-called containment rings. Containment rings, generally of wrought metal, are disposed radially outwardly of the turbine wheel and to some degree, axially to each side of the turbine wheel. The containment rings are made of material with sufficient strength that upon catastrophic failure of a turbine wheel, all parts thereof moving radially outwardly will impinge upon the containment ring which in turn will halt and arrest any further radial outward movement thereof.
While strength is the main constraint in providing a containment ring for land based or marine gas turbines, in the case of airborne gas turbines, weight becomes a constraint as well. And of course, cost is always of concern.
It can be readily appreciated that the greater the radial spacing from the axis of rotation of the turbine wheel to the containment ring, the greater the mass of the turbine ring because its circumference will increase in relation to the radial distance. Thus, it is desirable in aircraft gas turbines to maintain the containment ring as close as possible to the turbine wheel rotational axis. This also reduces the cost of materials where the cost of the material used in fabricating the containment ring is a significant factor.
However, in so doing, the containment ring is brought closer and closer to those areas of the gas turbine that are exposed to high temperatures; and this in turn means that increasingly exotic material whose cost is a consideration must be used in order to withstand the high temperatures and thermal cycling in the environment in which they are placed and yet reliably provide containment. Consequently, high cost, wrought containment ring structures have been employed as a trade off to obtain minimal mass.
It has also long been known in gas turbine engines to provide a dilution air zone in the combustor. This zone is conventionally located directly within the combustion annulus downstream of the fuel injectors but well upstream of the outlet of the combustor. Generally speaking, dilution air is injected into the combustion annulus to control the temperature of hot gases.
More particularly, upstream of the dilution zone both fuel and air are injected and ignited in the combustion annulus. It is also conventional for there to be a cooling air film introduced along the walls of the combustion annulus upstream of the dilution zone. Of course, the hot gases that result from combustion then pass toward turbine blades.
As is known, it is important to be able to control the temperature of the hot gases as they enter the nozzle on their way to the turbine blades. This has conventionally been handled by means of a dilution zone within the combustion annulus well upstream of the outlet of the combustor in order to ensure mixing and cooling prior to entry into the nozzle. While effective, this means of controlling the temperature of the hot gases is not satisfactory in all respects.
More particularly, the need to provide a dilution zone in a combustion annulus upstream of the outlet of the combustor dictates the geometry. In other words, the length of the turbine has been controlled to a degree by the necessity of having a distinct dilution zone within the combustion annulus, i.e., there was no opportunity for shortening the length of the combustor in order to reduce weight and expense. However, conventional designs have also failed to address still another serious problem.
More specifically, the dilution air flow path is known to cool only a portion of the walls of the combustor. Thus, in a conventional annular combustor of a gas turbine, not only is it true that not all portions of the walls of the combustor are cooled, but the point of injection into the dilution zone has rendered it impossible to effect any significant cooling of the turbine shroud and, thus, of the nozzle and turbine blades. As a result, it has remained to provide a low cost, simple, reliable turbine shroud cooling.
As will be appreciated, these problems lead to adverse consequences on performance and life span. In other words, due to the heretofore recognized inability to provide an ultra-short combustor and a well-cooled turbine shroud, it has been impossible to achieve the highest levels of power and fuel economy as well as longer life for the various components such as the nozzle blade, turbine shroud, turbine blades, turbine exhaust duct, etc. Furthermore, if an ultra-short combustor could be provided, there would be less exhaust noise thereby reducing silencing problems.
In the previously identified copending application, a means of achieving such advantageous cooling is disclosed. According to the present invention, such advantageous cooling means are also employed to overcome the above-stated problem incurred in fabricating gas turbine containment rings of low mass.